US 6935117 B2
A gas turbine engine is piloted with a pilot flow of fuel delivered to a combustor as a liquid. A first additional flow of the fuel is also delivered to the combustor as a liquid. A second additional flow of the fuel is vaporized and delivered to the combustor as a vapor. A fuel injector may have passageways associated with each of the three flows.
1. A fuel injector for a gas turbine engine comprising:
a mounting flange;
a stem extending from a proximal portion at the mounting flange to a distal portion;
a nozzle proximate the stem distal portion;
a first passageway along and through the stem and extending from a first inlet to a first outlet at the nozzle, the first outlet comprising a first plurality of apertures injecting gasified fuel;
a second passageway along and through the stem and extending from a second inlet to a second outlet at the nozzle, the second outlet comprising a second plurality of apertures injecting liquid fuel, generally radially inboard of the first plurality of apertures; and
a third passageway along and through the stem and extending from a third inlet to a third outlet at the nozzle, the third outlet comprising at least one third aperture injecting liquid fuel, generally radially inboard of the first plurality of apertures;
said first, second, and third passageways in fluid communication with one or more liquid fuel tanks;
said gas turbine engine being mounted on an aircraft to produce thrust.
2. The apparatus of
the first passageway has an effective cross-sectional area larger than an effective cross-sectional area of the second passageway, and
the effective cross-sectional area of the first passageway is larger than an effective cross-sectional area of the third passageway.
3. The apparatus of
along major portions of respective lengths, the first, second, and third passageways are within respective first, second and third conduits.
4. The apparatus of
the first passageway includes an outlet plenum.
5. The apparatus of
the first, second, and third outlets are concentric.
6. The apparatus of
the first and second outlets are respective first and second circular arrays of outlet apertures.
7. The apparatus of
the third outlet is a single outlet aperture.
8. The apparatus of
an upstream portion of each of the first, second, and third passageways is formed by an upstream portion of a conduit protruding form an outboard surface of the mounting flange.
9. The apparatus of
10. The apparatus of
a heat shield having leg and foot portions.
11. The apparatus of
within the leg potion of the heat shield, a plurality of collar plates each have first, second, and third apertures accommodating leg portions of the first, second, and third conduits, respectively and an outer periphery in facing proximity to an interior surface of the heat shield leg portion.
12. The apparatus of
the first, second, and third conduits each have a foot portion;
the second and third conduit foot portions are held spaced apart by spacers secured to one of the two so as to permit differential thermal expansion; and
the third conduit foot portion and heat shield foot portion are held spaced apart by spacers secured to one of the two so as to permit differential thermal expansion.
The invention was made with U.S. Government support under contract F33615-95-C-2503 awarded by the United States Air Force. The U.S. Government has certain rights in the invention.
(1) Field of the Invention
The invention relates to gas turbine engine combustion. More particularly, the invention relates to fuel injection systems for aircraft gas turbine engines.
(2) Description of the Related Art
Common gas turbine engines are liquid fueled. In a typical arrangement, the engine's combustor has one or more fuel injectors, each of which has a main passageway with multiple outlets for introducing a main flow of fuel and a pilot passageway for introducing a pilot flow of fuel. The pilot flow is initiated to start the engine and may remain on throughout the engine's operating envelope. The main flow may be initialized only above idle conditions and may be modulated to control the engine's output (e.g., thrust for an aircraft). For variety of performance reasons, it is known to use gaseous fuel (including a vaporized liquid). It is also known to use fuel as a heatsink.
Accordingly, one aspect of the invention involves a fuel injector for a gas turbine engine. The injector includes a mounting flange, a stem extending from a proximal portion at the mounting flange to a distal portion, and a nozzle proximate the stem distal portion. A first passageway extends through the stem from a first inlet to a first outlet at the nozzle. The first outlet has a number of apertures. A second passageway extends through the stem from a second inlet to a second outlet at the nozzle. The second outlet comprises a number of apertures, generally inboard of the apertures of the first passageway. A third passageway extends through the stem from a third inlet to a third outlet at the nozzle. The third outlet has at least one aperture generally inboard of the apertures of the first passageway.
In various implementations, the first passageway may have an affective cross-sectional area larger than an affective cross-sectional area of the second passageway. The affective cross-sectional area of the first passageway may be larger than an affective cross-sectional area of the third passageway. Along major portions of respective lengths, the first, second, and third passageways may be within respective first, second, and third conduits. The first passageway may include an outlet plenum.
Another aspect of the invention involves a combustor system for a gas turbine engine. A combustion chamber has at least one air inlet for receiving air. There is at least a first source of a gaseous first fuel and at least a second source of an essentially liquid second fuel. At least one fuel injector is positioned to introduce the first and second fuels to the air.
In various implementations, the first and second sources may comprise portions of a fuel system having a liquid fuel supply common to the first and second sources, with the second source vaporizing the liquid fuel to form the first fuel. The injectors may have a pilot passageway for carrying a pilot portion of the second fuel, a main liquid passageway for carrying a second portion of the second fuel, and a gaseous fuel passageway for carrying the first fuel.
Another aspect of the invention involves a method for fueling a gas turbine engine associated with a source of fuel in liquid form. The engine is piloted with a pilot flow of the fuel delivered to a combustor as a liquid. A first additional flow of the fuel is also delivered to the combustor as a liquid. A portion of the fuel is vaporized and delivered as a second additional flow of the fuel to the combustor as a vapor.
In various implementations, in at least certain conditions the first and second additional flows may be simultaneous. A mass flow of the second additional flow may be 40–70% of a total main burner fuel flow. The vaporizing may comprise drawing heat to the portion from at least one system on or associated with the engine. A ratio of the first flow to the second flow may be dynamically balanced based upon a combination desired heat extraction from the at least one system and a desired total fuel flow for the engine.
The details of one or more embodiments of the invention are set forth in the accompanying drawings and the description below. Other features, objects, and advantages of the invention will be apparent from the description and drawings, and from the claims.
Like reference numbers and designations in the various drawings indicate like elements.
The exemplary fuel injector 40 (
Along the injector foot, the foot portion 110 of the first conduit 60 passes through an aperture 112 in the second conduit 62 near the intersection of the leg and plenum portions of the second passageway. There the first conduit is secured to the second conduit such as by brazing. Similarly, an end portion of the first conduit 60 may be secured within an aperture 114 in the end plate 106. This securing is appropriate as there is relatively little stress between the first and second conduits when both are carrying liquid fuel. However, the inner wall 92 of the foot portion of the third conduit is held spaced-apart from the outer wall 104 of the foot portion of the second conduit by spacers 120. Advantageously, the spacers may float with respect to one of these two conduits and be secured to the other. This permits relatively free floating differential thermal expansion of the third conduit relative to the second and first as the former may be more highly heated by the gaseous fuel it carries.
Externally, the injector includes a heat shield having leg and foot portions 130 and 132. As with the second and third conduit foot portions, the third conduit foot portion and heat shield foot portion are held spaced apart by spacers 134 which may be secured to one of the two so as to permit differential thermal expansion. Within the leg, there may be several collar plates 140 having three apertures for accommodating the leg portions of the three conduits and an outer periphery 142 (
Additionally, there are one or more flow paths 180 for delivering fuel as a gas. The gas and liquid flow paths may partially overlap and, within either family, the flow paths may partially overlap. The gaseous flow paths include heat exchangers 182 for transferring heat to liquid fuel along such gaseous flow paths to vaporize such fuel. In the exemplary embodiment, the heat exchangers are fluid-to-fluid heat exchanges for drawing heat from one or more heat donor fluids flowing along one or more fluid flow paths 190. Exemplary heat donor fluid is air from the high pressure compressor exit. Gaseous fuel delivery is governed by one or more pressure regulating valves 192 downstream of the heat exchangers. Control valves 194 in the donor flow paths may provide control over the amount of flow through such donor flow paths.
In operation, the desired engine output will essentially determine the desired total amount of fuel. The desired heat extraction from the donor flow path 190 will essentially determine the amount of such fuel which passes along the gaseous flow paths 180. Although the temperatures of the liquid fuel in the reservoir and of the discharge vapor may vary, the latent heat of vaporization strongly ties the mass flow rate of vaporized fuel to the desired heat extraction. In operation, therefore, the control system (not shown) may dynamically balance the proportions of fuel delivered as liquid and delivered as vapor in view of the desired heat transfer. In operation, mass flow rates of the pilot fuel relative to the total may be small (e.g., less than 10% for the pilot fuel at subsonic cruise conditions). The high pressure compressor experiences high temperatures generated at high flight Mach numbers. Thus, greater cruise heat transfer will be required at supersonic conditions, biasing a desirable balance toward vapor at such speeds. The system may be sized such that the main liquid fuel flow reaches a capacity limit at an intermediate power. Thus at higher power non-cruise conditions (e.g., up to max. power), both heat transfer and high total fuel requirements may indicate substantial use of the vaporized fuel in addition to a maximal flow of liquid fuel, thus also biasing toward vapor (at least relative to a low or zero vapor flow at low subsonic cruise conditions).
In one example, at maximum dry power operation the vapor system could be employed at Mach numbers greater than 0.5, whereas at cruise or part power operation the vapor system could be employed at Mach numbers greater than 1.0. The mass flow rate of fuel delivered along the third flow path may be 40–70% of a total main burner (e.g., exclusive of augmentor) fuel flow at an exemplary supersonic cruise condition, 30–50% at an exemplary subsonic cruise condition, 40–70% at an exemplary subsonic max power condition, and 60–80% at an exemplary supersonic max. power condition. A ratio of the effective cross-sectional areas of the second and third passageways may be between 1:2 and 1:4.
One or more embodiments of the present invention have been described. Nevertheless, it will be understood that various modifications may be made without departing from the spirit and scope of the invention. For example, the invention may be applied to a variety of existing or other combustion system configurations. The details of such underlying configurations may influence details of any particular implementation. Accordingly, other embodiments are within the scope of the following claims.